1. Field of the Invention
The present invention relates to small, compact and expendable gas turbine engine having a reduced cost of construction and improved performance and operating life in order to increase the range of operation.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Small gas turbine engines of the kind used in Unmanned Air Vehicles (UAV) such as a small cruise missile or a drone are well known in the art. These turbines produce a thrust from less than 300 lbs to several hundred lbs. Because these turbine engines must fit within a small space, they tend to be very compact. Since the engine must be compact in size, the combustor must be as small as possible. However, the combustor must provide a long enough burn path to remain lit, and to burn the fuel to produce power without wasting unburned fuel. In small combustors, the fuel droplets size must be small in order to burn in the smaller combustor sections in which the fuel particle residence time for burning the particles is short. Larger fuel particles will take longer to burn and in a small combustor will not burn completely. Effective us of the available volume must be made so that the combustor is able to provide the required heat output over a range of engine operating speeds and flight conditions.
Small expendable gas turbine engines also suffer from poor specific fuel consumption (SFC), which limits the engine to unnecessarily short range and loiter times. Small engines suffer from restricted flight/starting envelopes and operating speeds due to combustor flame-out limitations.
U.S. Pat. No. 3,381,471 issued to Szydlowski on May 7, 1968 shows a combustion chamber for gas turbine engines in which the combustion chamber includes a combustion chamber space 24 and a dilution space 25, where an injection wheel 27 injects fuel into the combustion space 24 for burning, and additional air is added in the dilution space to complete the burning of the fuel that didn't burn in the combustion space.
Another Prior Art swirl combustor is shown in the U.S. Pat. No. 4,996,838 issued to Melconian on Mar. 5, 1991 entitled annular vortex slinger combustor. Melconian shows a combustor having a primary annular combustion chamber 18 and a secondary annular combustion chamber 20, where fuel is injected into the primary combustion chamber through fuel injectors 24 located in the forward portion of the primary combustion chamber or in a different embodiment by a fuel nozzle 100 rotating about the primary axis to deliver fuel into the primary combustion chamber. Louvers 22 are peripherally disposed circumferentially about the inner and outer walls of the primary and secondary chambers to deliver compressed air in a helical direction into both combustion chambers. The combustion of unburned gaseous products from the primary combustion chamber is completed in the secondary combustion chamber.
U.S. Pat. No. 4,040,251 issued to Heitmann et al. on Aug. 9, 1977 entitled gas turbine combustor chamber arrangement in which a combustor includes a primary zone 60 and a dilution zone 62 located downstream from the primary zone. Fuel is injected into the primary zone by a fuel slinger 34 mounted on a shaft 36 which is coupled to rotate with the compressor impeller.
In the Szydlowski, Melconian, and Heitmann inventions discussed above, all three of these combustors suffer from the same problem: combustion space is large in order to provide the required space to allow for enough fuel to burn to generate the gas stream for the engine. Therefore, the combustor would not be practical in a small, compact gas turbine engine used in a UAV. The addition of the secondary combustor section downstream of the primary combustor section does not reduce the overall size of the combustor to a size that would be practical in today's UAV which is smaller than a cruise missile.
U.S. Pat. No. 5,526,640 issued to Brooks et al on Jun. 18, 1996 shows a gas turbine engine with a primary combustion zone 40a and a secondary combustion zone 40b, a forward main bearing 46 and a rear main bearing 48, a fuel slinger 38 to inject fuel into the primary combustion zone. Air flows from a compressor into an outer air annulus 60, through air tubes 64, into an inner air annulus 62, through an opening 36d and into an air/fuel annulus 96, into a slinger 38, and then into the primary combustion zone 40a. The air/fuel mixture passing through the annulus flow path 96 is used to lubricate both main bearings 46 and 48 and to cool the turbine hub portion 30a. The second embodiment of FIG. 9 shows similar structure. However, Brooks does not show a fuel injector for the secondary combustion zone. Also, the flow rate of air through the rear main bearing is in the range of 2% or less of the total air flow rate into the combustor. This very low flow rate would not provide enough cooling for the bearing.
U.S. Pat. No. 5,727,378 issued to Seymour on Mar. 17, 1998 entitled Gas Turbine Engine with a main bearing located near a turbine section, where a portion of the compressor air is diverted into a cooling cavity 49 that cools the bearing near the turbine wheel, and then into the turbine section without passing through the combustor. Both main bearings are located upstream of the combustor and the turbine sections of the engine. In the Seymour invention, the combustor is not positioned between the two main bearings, and the cooling air for the one main bearing is not used in the combustor.
It is an object of the present invention to provide for a gas turbine engine small enough to fit within a small UAV such as a small cruise missile or drone.
It is another object of the present invention to provide for a small gas turbine engine that has improved operating times and low weight in order to increase the range and loiter time of the UAV.
An additional object of the present invention is to make a low cost and easier to manufacture gas turbine engine having a low parts counts with a minimal number of fuel injection points.
It is another object of the present invention to provide a small gas turbine engine with a rotary cup combustor to deliver fuel and air to a primary burn zone and a secondary burn zone in the combustor in order to maximize combustor operating range and efficiency as well as to provide for a low pressure source to draw the cooling air across the bearing for cooling purposes. The rotary cup combustor will allow for heavy fuels to be burned in that the rotary cup combustor provides excellent fuel break-up and large fuel passage sizes.
Another object of the present invention is to use the rotary cup as a fuel pump by passing the fuel through the rotary shaft and into the rotary cup, where the fuel is pumped to high pressure due to rotation of the shaft, allowing for the fuel pump to be significantly smaller, lower cost and more precise with respect to fuel metering.
Another object of the present invention is to provide for a single igniter at the outer combustor wall, the igniter being located in a stagnation point between the primary burn zone and the secondary burn zone, in which the fuel is sprayed directly toward the igniter where the flow stagnation point is located, providing for the widest possible engine relight envelope.
These objects and others will be described below in the detailed description of the invention and the accompanying drawings.